Field of the Invention
The invention relates to a method for combustion of a fuel in a flow of compressed air which passes through a gas turbine from a compressor section to a turbine section, wherein the fuel is added to the flow in the compressor section and is burnt between the compressor section and the turbine section. The invention also relates to a corresponding gas turbine.
Such a method and such a gas turbine have been disclosed in U.S. Pat. No. 2,630,678.
Published European Patent Application 0 590 297 A1 discloses a gas turbine having a compressor section, an annular combustion chamber and a turbine section. The compressor section provides a flow of compressed air which has fuel added to it in the annular combustion chamber after which the fuel is ignited and burnt. The flow is passed to the turbine section after the combustion has taken place. That document refers to the gas turbine as a "gas turbine assembly", the compressor section as a "compressor" and the turbine section as a "turbine". The different terminology is a result of the fact that the term "gas turbine" is not used in a standard manner in the specialist world. The term "gas turbine" may refer both to a turbine in the narrow sense, that is to say an engine which extracts mechanical energy from a flow of heated gas, and to a unit including a turbine in the narrow sense as well as a combustion chamber or combustion chambers and a compressor section. In the present context, the term "gas turbine" always refers to a unit which, in addition to a turbine in the narrow sense, that is always referred to as a "turbine section" in this document, also includes at least one associated compressor section.
Examples of burners which can be used in a gas turbine can be found in Published European Patent Application 0 193 838 B1, U.S. Pat. No. Re. 33896, Published European Patent Application 0 276 696 B1 and U.S. Pat. No. 5,062,792. A combustion chamber in the form of an annular combustion chamber having a multiplicity of burners disposed in the form of an annular ring is described in Published European Patent Application 0 489 193 A1.
Further information relating to the construction of a combustion device which can be disposed between a compressor section and a turbine section of a gas turbine is disclosed in U.S. Pat. Nos. 2,755,623; 3,019,606; 3,701,255 and 5,207,064. That information includes configurations for the implementation of combustion devices in which a flow of compressed air is carried with a spin and the combustion possibly also takes place in the spinning flow. Those documents also contain information about components, in particular about flame holders, which are intended to stabilize a combustion process.
One important source of thermodynamic losses is a pressure loss which occurs between the compressor section and the turbine section, that is to say over that region of the gas turbine where the flow of compressed air is heated by combustion of a fuel. That pressure loss is governed by the high level of structural complexity, which has always been accepted until now, to produce a combustion device in the form of one or more combustion chambers. Certain rules for reducing the complexity are known. In particular, the already mentioned Published European Patent Application 0 590 297 A1 discloses a so-called "annular combustion chamber" in which the flow is intended to maintain a spin, to which it is subjected in the compressor section, during the combustion of the fuel so that there is no need for any conventional stationary ring of blades at an inlet to the turbine section, in order to initially build up any spin required to operate the turbine section. Reference is also made to U.S. Pat. No. 2,630,678, which was cited initially, and according to which the fuel can be added in the compressor section itself.
In addition to the already mentioned measures for improving the thermodynamic process which takes place in the gas turbine, the increase in the specific power, that is to say the power emitted by the gas turbine per unit amount of energy supplied with the fuel, necessitates an increase in the turbine inlet temperature, that is to say the temperature of the flow after combustion of the fuel and upon entry into the turbine section. The turbine inlet temperature is limited by the load capacity of the components in the turbine section, which is governed in particular by the load capacity of the materials being used and the measures which may be provided to cool the components. Such measures are normally limited by the fact that air required for cooling must be tapped off the flow and is no longer available for combustion. The distribution of the temperature in the flow upon entry into the turbine section is also important. If the distribution of the temperature in the flow upon entry into the turbine section is not uniform, as must be assumed for every turbine produced to date, then the maximum temperature in the flow governs the maximum load on the components in the turbine section and, in order to operate the latter safely, therefore has to be kept below a critical limit while, in contrast, the mean value of the temperature in the flow is the governing factor for the quality of the thermodynamic process and, in particular, for that mechanical power which the thermodynamic process can provide for a given use of primary energy. It follows from those considerations that the specific power of a gas turbine can be increased, without any adverse effect on its life, if it is possible to homogenize the distribution of the temperature in the flow upon entry into the turbine section, and thus to raise the mean value of the temperature to the maximum temperature. Once homogenization has been carried out, the mean value of the temperature in the flow can be raised by increasing the use of primary energy until the predetermined load capacity of the turbine section is reached. The potential of such measures is considerable. Raising the mean value of the temperature in the flow upon entry into the turbine section by about 10.degree. C. can produce an increase in the specific power of more than 1%. Conventional gas turbines invariably have the potential for such measures since the difference between the maximum and the mean value in the distribution of the temperature in the air flow upon entry into a turbine section in such gas turbines is up to 100.degree. C.
The reason for the inhomogeneous distribution of temperature in a flow in a conventional gas turbine is normally the complex and inherently inhomogeneous treatment of the flow and of the fuel between the compressor section and the turbine section. That is true to a particular extent if the flow is split into flow elements and is fed to a plurality of combustion chambers or to a plurality of individual burners.
That is also true in conventional annular combustion chambers, which in each case largely dispense with any splitting of the flow but still provide a plurality of burners, that are necessary at a distance from one another and are intended to heat the flow.
Furthermore, it is necessary to take account of the fact that, in any conventional gas turbine, the flow of compressed air between the compressor section and the turbine section, that is to say where it is heated by combustion of a fuel, is carried without any spin. The major reason therefor is that such a measure can reduce the speed of the flow to a minimum. That is the easiest way to ensure stable combustion of the fuel, while providing maximum flexibility for the construction of burners and the like. In fact, conventional practice demands that guidance devices be provided at the end of the compressor section which extract from the flow any spin that exists downstream of the last rotating compressor stage and, in addition, the turbine section has to have a guidance device at its inlet, which provides the flow with a spin required to act on the first rotating turbine stage. The guidance device in the turbine section, in particular, is the most severely thermally loaded component and must have a correspondingly complex construction. In addition, some pressure reduction occurs even in that guidance device, and thus a temperature reduction, of the combustion gas in the flow. Accordingly, it is not the first rotating turbine stage that governs the maximum possible temperature of the flow, but the guidance device at the inlet of the turbine section which, in fact, does not extract any energy from the flow.
The considerations discussed in the last two paragraphs are of particular importance for modern gas turbines, which are always characterized by the fact that they largely make full use of the limits predetermined by the materials being used. That is done particularly to achieve the maximum possible thermodynamic efficiencies. Gas turbines for stationary use, which have ratings of between 100 MW and 250 MW, have compressor sections which are characterized by pressure ratios between 16 and 30, corresponding to temperatures of between 400.degree. C. and 550.degree. C. at the respective compressor outlet, and as a result of the combustion provide heated combustion gas which reaches temperatures of between 1100.degree. C. and 1400.degree. C. All of the temperatures require the greatest possible care in the construction of the combustion devices and turbine sections and full utilization of the limits predetermined by the materials being used. In particular, the temperatures quoted for compressor outlets must also be regarded as being critical in terms of possible self-ignition of the fuel that is added.